1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled transition duct positioned between a combustor outlet and a turbine inlet.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a combustor that produces a hot gas flow and a turbine that receives the hot gas flow to produce mechanical work. The turbine includes a row of first stage stator vanes that receive the hot gas flow from the combustor and guide the flow into the first stage rotor blades. The highest temperature gas flow is found between the combustor outlet and the first stage rotor blades.
A combustor can be an annular combustor formed from two concentric walls or a plurality of can combustors arranged in an annular array around the engine. Annular combustors are typically used in aircraft engines. Can combustors are typically used in large industrial gas turbine engines. In both combustor types, a transition piece is needed to guide the hot gas flow from the combustor exit into the first stage guide vanes. In the annular combustor, the combustor exit is already annular in shape so not much redirecting is required. The exit flow from the annular combustor can flow directly into the first stage guide vanes.
In the can combustor arrangement, the hot gas flow from the combustor exit must be transformed from a circular flow to a segment of an annular flow. A transition duct in the IGT (industrial gas turbine) engine is circular in the inlet end and some-what rectangular in the outlet end so that the combined flow of all of the can combustors will be in an annular flow path at the inlet to the first stage stator vanes. The transition duct of the engine is exposed to the highest gas flow temperature. Thus, improved cooling of the transition duct will increase the part life.